Geared gas turbine engine

ABSTRACT

A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine , a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.

CROSS-REFERENCE TO RELATED APPLICATION(S

This application is a continuation of U.S. Application No. 17/411,617filed Aug. 25, 2021, which is a continuation of U.S. Application No.16/526,221 filed Jul. 30, 2019, which is based on and claims priorityunder 35 U.S.C. 119 from Great Britain Patent Application No. 1907256.0filed on May 23, 2019. The entire contents of the above applications areincorporated herein by reference.

The present disclosure relates to a gas turbine engine for an aircraftand a method of operating a gas turbine engine on an aircraft.

Turbofan gas turbine engines for aircraft propulsion have many designfactors that affect the overall efficiency and power output or thrust.To enable a higher thrust while maintaining efficiency, a largerdiameter fan may be used. As the diameter of the fan is increased,however, the required lower speed of the fan tends to conflict with therequirements of the turbine component the core shaft is connected to,typically a low pressure turbine. A more optimal combination can beachieved by including a gearbox between the fan and the core shaft,which allows the fan to operate at a reduced rotational speed, andtherefore enables a larger size fan, while maintaining a high rotationalspeed for the low pressure turbine, enabling the overall diameter of theturbine to be reduced.

A high propulsive efficiency for a geared gas turbine engine is achievedthrough a high mass flow through the engine. This may be enabled in partby increasing the bypass ratio of the engine, which is the ratio betweenthe mass flow rate of the bypass stream to the mass flow rate enteringthe engine core. To achieve a high bypass ratio with a larger fan whilemaintaining an optimum gearing ratio and fan speed, the size of theengine core, in particular the low pressure turbine, may need toincrease, which would make integration of a larger fan engine underneathan aircraft wing more difficult. A general problem to be addressedtherefore is how to achieve a high propulsive efficiency for a largergeared gas turbine engine while enabling the engine to be integratedwith an aircraft.

According to a first aspect there is provided a gas turbine engine foran aircraft comprising:

-   an engine core comprising a turbine, a compressor, and a core shaft    connecting the turbine to the compressor;-   a fan located upstream of the engine core, the fan comprising a    plurality of fan blades;-   a nacelle surrounding the engine core and defining a bypass duct and    bypass exhaust nozzle; and-   a gearbox that receives an input from the core shaft and outputs    drive to the fan so as to drive the fan at a lower rotational speed    than the core shaft,-   wherein the gas turbine engine is configured such that a jet    velocity ratio of a first jet velocity exiting from the bypass    exhaust nozzle to a second jet velocity exiting from an exhaust    nozzle of the engine core at idle conditions is greater by a factor    of around 2 or more than the jet velocity ratio at maximum take-off    conditions.

A large variation in the jet velocity ratio between idle and maximumthrust enables the core and bypass streams to be managed to keep theengine operable with a high propulsive efficiency and a high bypassratio. The large variation may be achieved by optimising the aerodynamicdesign of the fan and compressor components and/or by using otherdevices on the engine such as variable guide vanes and bleeds.

The jet velocity ratio, R_(J), may be defined as:

$R_{J} = \frac{V_{B}C_{B}}{V_{C}C_{C}\eta_{LPT}\eta_{F}}$

where V_(B) is the fully expanded first jet velocity, C_(B) is a thrustcoefficient of the bypass nozzle, V_(C) is the fully expanded second jetvelocity, C_(C) is a thrust coefficient of the core exhaust nozzle,η_(LPT) is an isentropic efficiency of a lowest pressure turbine of theengine core and η_(F) is an isentropic efficiency of compression of airinto the bypass duct by the fan. The fully expanded jet velocity may bedefined as the axial jet velocity at the point where the exhaust jet hasexpanded to ambient pressure.

The gearbox may have a gear ratio of between around 2.5 and around 5, ormay have a gear ratio within a range as defined in more detail below.

In some examples the gas turbine engine may be configured such that thejet velocity ratio is within a range from around 0.75 to around 1.3 atcruise conditions.

The gas turbine engine may be configured such that the jet velocityratio at idle conditions is between around 2 and 3.

The gas turbine engine may be configured such that the jet velocityratio at maximum take-off conditions is between around 0.75 and 1.3, oroptionally between around 0.8 and 1.0.

The fan may have an outer diameter of between around 220 cm and around290 cm, and optionally between around 230 cm and around 260 cm, or mayhave an outer diameter within a range as defined below.

The factor relating to the difference in jet velocity ratio between idleand maximum take-off conditions may be defined as

$\frac{Rj\,\, at\,\, Ground\, Idle}{Rj\,\, at\,\, Take - Off},$

i.e. the jet velocity ratio at ground idle conditions divided by the jetvelocity ratio at maximum take-off conditions.

The factor may be within a range of between around 2 and around 3.5, oroptionally between around 2.1 and around 3.16. Above the lower limit ofaround 2 or 2.1 enables a reduced fuel burn and can be achieved byfeatures such as a straighter fan root to maintain fan operability at alower specific thrust achieved by a larger fan diameter in combinationwith a gearbox and a smaller engine core. Above the upper limit ofaround 3.16 or 3.5, the fan diameter required becomes increasinglyunacceptable for a below-wing installation and would require furtherfeatures to reduce drag.

Maximum take-off (MTO) conditions may be defined as operating the engineat International Standard Atmosphere (ISA) sea level pressure andtemperature conditions +15° C. at maximum take-off thrust at end ofrunway, which is typically defined at an aircraft speed of around 0.25Mn, or between around 0.24 and 0.27 Mn. Maximum take-off conditions forthe engine may therefore be defined as operating the engine at a maximumtake-off thrust at ISA sea level pressure and temperature +15° C. with afan inlet velocity of 0.25 Mn.

Idle conditions may be defined as operating the engine at around 4% ofmaximum take-off thrust at ISA sea level pressure and temperatureconditions +15° C. Alternatively, idle conditions may be defined asoperating the engine at a minimum steady state thrust under ISA sealevel pressure and temperature conditions +15° C.

According to a second aspect there is provided a method of operating agas turbine engine on an aircraft, the gas turbine engine comprising:

-   an engine core comprising a turbine, a compressor, and a core shaft    connecting the turbine to the compressor;-   a fan located upstream of the engine core, the fan comprising a    plurality of fan blades; and-   a gearbox that receives an input from the core shaft to drive the    fan at a lower rotational speed than the core shaft,-   wherein the method comprises operating the gas turbine engine to    provide propulsion such that a jet velocity ratio between a first    jet velocity exiting from a bypass duct of the engine and a second    jet velocity exiting from an exhaust nozzle of the engine core    varies by more than a factor of two between idle and maximum thrust    at ISA sea level conditions.

The various optional and advantageous features associated with theinvention according to the first aspect may also apply to the secondaspect.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the The gearbox may be a reduction gearbox (in that theoutput to the fan is a lower rotational rate than the input from thecore shaft). Any type of gearbox may be used. For example, the gearboxmay be a “planetary” or “star” gearbox, as described in more detailelsewhere herein. The gearbox may have any desired reduction ratio(defined as the rotational speed of the input shaft divided by therotational speed of the output shaft), for example greater than 2.5, forexample in the range of from 3 to 4.2, or 3.2 to 3.8, for example on theorder of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4,4.1 or 4.2. The gear ratio may be, for example, between any two of thevalues in the previous sentence. Purely by way of example, the gearboxmay be a “star” gearbox having a ratio in the range of from 3.1 or 3.2to 3.8. In some arrangements, the gear ratio may be outside theseranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable).

The row of rotor blades and the row of stator vanes may be axiallyoffset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32. These ratios maycommonly be referred to as the hub-to-tip ratio. The radius at the huband the radius at the tip may both be measured at the leading edge (oraxially forwardmost) part of the blade. The hub-to-tip ratio refers, ofcourse, to the gas-washed portion of the fan blade, i.e. the portionradially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches),260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm(around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches ),350 cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380(around 150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm(around 160 inches) or 420 cm (around 165 inches). The fan diameter maybe in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 320 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tiP) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tiP) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (allunits in this paragraph being Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading maybe in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of from 0.28 to 0.31 or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratiomay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of from 13 to 16, or 13 to 15, or 13 to 14. Thebypass duct may be substantially annular. The bypass duct may beradially outside the core engine. The radially outer surface of thebypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 80 Nkg⁻¹s to100 Nkg⁻¹s, or 85 Nkg⁻¹s to 95 Nkg⁻¹s. Such engines may be particularlyefficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15° C. (ambient pressure 101.3 kPa, temperature 30° C.), with the enginestatic.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400 K, 1450 K, 1500 K,1550 K, 1600 K or 1650 K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K or 2000 K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 1800 K to 1950 K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

Maximum take-off thrust for the engine may be defined as operating theengine within 15° C. of International Standard Atmosphere sea levelpressure and temperature conditions at maximum take-off thrust at end ofrunway, which is typically defined at an aircraft speed of around 0.25Mn, or between around 0.24 and 0.27 Mn. Maximum take-off conditions forthe engine may therefore be defined as operating the engine at a maximumtake-off thrust at ISA sea level pressure and temperature with a faninlet velocity of 0.25 Mn.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint – in terms of time and/or distance–between top of climb and start of descent. Cruise conditions thus definean operating point of the gas turbine engine that provides a thrust thatwould ensure steady state operation (i.e. maintaining a constantaltitude and constant Mach Number) at mid-cruise of an aircraft to whichit is designed to be attached, taking into account the number of enginesprovided to that aircraft. For example where an engine is designed to beattached to an aircraft that has two engines of the same type, at cruiseconditions the engine provides half of the total thrust that would berequired for steady state operation of that aircraft at mid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide – in combination withany other engines on the aircraft – steady state operation of theaircraft to which it is designed to be attached at a given mid-cruiseMach Number) at the mid-cruise atmospheric conditions (defined by theInternational Standard Atmosphere according to ISO 2533 at themid-cruise altitude). For any given gas turbine engine for an aircraft,the mid-cruise thrust, atmospheric conditions and Mach Number are known,and thus the operating point of the engine at cruise conditions isclearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein, and/or the maximum take-offconditions relate to the maximum take-off conditions of the aircraft.

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise and/or maximum take-off of the aircraft, asdefined elsewhere herein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is an example plot of change in fuel burn as a function of jetvelocity ratio;

FIG. 5 is a schematic drawing of an aircraft having a gas turbine enginemounted thereon; and

FIG. 6 is a schematic drawing illustrating the concept of a fullyexpanded jet velocity.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2 . The low pressure turbine 19 (see FIG. 1 ) drives the shaft26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclicgear arrangement 30. Radially outwardly of the sun gear 28 andintermeshing therewith is a plurality of planet gears 32 that arecoupled together by a planet carrier 34. The planet carrier 34constrains the planet gears 32 to precess around the sun gear 28 insynchronicity whilst enabling each planet gear 32 to rotate about itsown axis. The planet carrier 34 is coupled via linkages 36 to the fan 23in order to drive its rotation about the engine axis 9. Radiallyoutwardly of the planet gears 32 and intermeshing therewith is anannulus or ring gear 38 that is coupled, via linkages 40, to astationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3 . Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3 . There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2 . For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2 .

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1 ), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

FIG. 4 illustrates an example plot of change in fuel burn, ΔFB, as afunction of jet velocity ratio, R_(J)., other factors being constant.The change in fuel burn contribution from propulsive efficiency 401 isdetermined relative to an optimum value for the jet velocity ratio ofaround 1.0, with an increase in fuel burn both above and below thisvalue. Factors that may affect the jet velocity ratio include therelative rotational speeds of the fan and turbine and the areas of theexhaust nozzles for the bypass and core exhausts. A lower gear ratio ofthe gearbox, i.e. a gear ratio of around 3.4 or lower, will tend toresult in values for the jet velocity ratio of 1.0 or greater. To keepthe fuel burn loss to within around 0.5% or less, it can be seen fromFIG. 4 that the jet velocity ratio should be between around 1.0 andaround 1.3. As the jet velocity ratio increases, the increase in fuelburn becomes greater. A further preferred upper limit for the jetvelocity ratio is around 1.2, which keeps the increase in fuel burn toaround 0.25-0.3%.

For a higher gearing ratio, i.e. around 3.3 to 3.4 and above, forexample up to around 3.8 or in some cases even higher, the jet velocityratio tends to be around 1.0 or below. As the jet velocity ratiodecreases, the fuel burn contribution from propulsive efficiency 401increases, and at a higher rate than for the portion above 1.0. Tomaintain this loss to within around 0.5%, it can be seen from FIG. 4that the jet velocity ratio should be kept within around 0.8 to around1.0, and for a ratio of around 0.75 and below, the fuel burncontribution from propulsive efficiency becomes dominant, rising toaround 0.7% and above. A further preferred lower limit for the jetvelocity ratio of around 0.85 or 0.90 may be used to keep the fuel burncontribution from propulsive efficiency to around 0.25% or below.However, further decreasing the jet velocity ratio enables a higher gearratio to be used and/or decreases the pressure ratio across the IPturbine, thereby allowing for a smaller, faster and/or lighter IPturbine, reflected in a lower contribution to fuel burn loss 402 by theIP turbine. A range of around 0.75 to around 0.82 for the jet velocityratio is thereby advantageous.

For a given set of gears making up an epicyclic gearbox, a planetarydriving arrangement will produce a higher gearing ratio than a stardriving arrangement. A star arrangement is therefore generally preferredin combination with a jet velocity ratio of around 1.0 and above, and aplanetary arrangement for a jet velocity ratio of around 1.0 and below.

In a general aspect therefore, the gas turbine engine may be configuredsuch that the jet velocity ratio is within a range from around 0.75 toaround 1.3 at cruise conditions.

FIG. 5 illustrates an example aircraft 50 having a gas turbine engine 10attached to each wing 51 a, 51 b thereof. When the aircraft is flyingunder cruise conditions, as defined herein, each gas turbine engine 10operates such that a jet velocity ratio between a first jet velocityexiting from a bypass duct of the engine 10 and a second jet velocityexiting from an exhaust nozzle 20 of the engine core is within a rangefrom around 0.75 to around 1.3.

FIG. 6 illustrates an example exhaust nozzle 60 of a gas turbine engine.The pressure Pj at the exit or throat 61 of the exhaust nozzle 60 isgreater than the ambient pressure Pa around the engine. At some distanceaway from the nozzle exit 61 the jet pressure will equalise with theambient pressure, i.e. Pj=Pa. The fully expanded jet velocity is definedas the jet velocity 62 at this point, i.e. the jet velocity along theaxis of the engine at a minimum distance from the exhaust nozzle wherethe pressure is equal to ambient pressure.

Parameters that may be adjusted to achieve a jet velocity ratio withinthe desired range may include the LPT blade exit angle, LPT exit area,and the LPT rotation speed.

The following table illustrates example parameters for two engineexamples, example 1 being for a relatively small, or lower power, engineand example 2 for a relatively large, or higher power, engine. A smallengine may for example have a fan diameter of between around 200 and 280cm and/or a maximum net thrust of between around 160 and 250 kN or asdefined elsewhere herein. A large engine may for example have a fandiameter of between around 310 and 380 cm and/or a maximum net thrust ofbetween around 310 and 450 kN or as defined elsewhere herein.

Parameter Example 1 (small engine) Example 2 (large engine) Fan diameter(cm) 215 320 LPT Exit Total Pressure at maximum flow (kPa) 130 130Maximum LPT Exit Mass Flow (kg/s) 50 100 LPT Final Rotor Area (m²) 0.38or less, for example 0.25 to 0.38 0.75 or less, for example 0.5 to 0.75ESS Inlet Total Pressure at maximum flow (kPa) 140 140 ESS Inlet MassFlow (kg/s) 50 100 ESS Inlet Rotor Area (m²) 0.275 or greater, forexample 0.27-0.3 0.55 or greater, for example 0.55-0.6

The above parameters relating to LPT exit total pressure at maximumflow, maximum LPT exit mass flow and LPT final rotor area togetherdetermine the exit flow velocity of the LPT, i.e. the flow velocity atan exit of the engine core. The ESS inlet total pressure at maximumflow, maximum ESS inlet mass flow and ESS inlet rotor area togetherdetermine the velocity at the inlet of the engine core. The axialexhaust flow velocity from the bypass exhaust nozzle may be determined,at least in part, by the area of the bypass exhaust nozzle outlet.

In order to achieve a jet velocity ratio within the desired range, thefan may be provided with features such as a straighter fan root. Thecompressors, in particular the high pressure compressor, may be providedwith features to manage their operability to allow the compressors tooperate at a low power required to meet the defined ratios, which mayfor example include devices such as variable guide vanes. This changesthe flow incidence onto the blades and helps to maintain an operabilitymargin preventing the compressor from surging or stalling when operatingat lower speeds.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and subcombinations of one or morefeatures described herein.

1. A method of operating a gas turbine engine on an aircraft, the gasturbine engine comprising: an engine core comprising a turbine, acombustor, a compressor, and a core shaft connecting the turbine to thecompressor; a fan located upstream of the engine core, the fancomprising a plurality of fan blades; a nacelle surrounding the enginecore and defining a bypass duct and bypass exhaust nozzle; and a gearboxthat receives an input from the core shaft (26) and outputs drive to thefan so as to drive the fan at a lower rotational speed than the coreshaft, wherein the method comprises operating the gas turbine engine toprovide propulsion such that a jet velocity ratio, R_(J), of a first jetvelocity exiting from the bypass exhaust nozzle to a second jet velocityexiting from an exhaust nozzle of the engine core is defined as:$R_{J} = \frac{V_{B}C_{B}}{V_{C}C_{C}\eta_{LPT}\eta_{F}}$ where V_(B) isa fully expanded first jet velocity, C_(B) is a thrust coefficient ofthe bypass exhaust nozzle, V_(C) is a fully expanded second jetvelocity, C_(C) is a thrust coefficient of the engine core exhaustnozzle, η_(LPT) is an isentropic efficiency of a lowest pressure turbineof the engine core and η_(F) is an isentropic efficiency of the fan tip;the jet velocity ratio, R_(J), is between around 0.75 and 1.3 at cruiseconditions; a fan tip loading defined as dH/U_(tip) ² is between 0.28and 0.35 at cruise conditions, where dH is the enthalpy rise across thefan and U_(tip) is the translational velocity of the leading edge of thefan tip; and the temperature of the flow at the exit of the combustor,at a position immediately upstream of a first turbine vane, is at least1600 K at cruise conditions.
 2. The method of claim 1, wherein thetemperature of the flow at the exit of the combustor, at a positionimmediately upstream of a first turbine vane, is between 1600 K and 1650K at cruise conditions.
 3. The method of claim 1, wherein thetemperature of the flow at the exit of the combustor, at a positionimmediately upstream of a first turbine vane, is between 1900 K and 2000K at maximum take-off conditions.
 4. The method of claim 1, wherein thefan tip loading at cruise conditions is between 0.28 and 0.33.
 5. Themethod of claim 1, wherein the fan tip loading at cruise conditions isbetween 0.28 and 0.30.
 6. The method of claim 1, wherein the jetvelocity ratio, R_(J), is between around 0.8 and 1.0 at maximum take-offconditions.
 7. The method of claim 1, wherein the jet velocity ratio,R_(J), is between around 2 and 3 at idle conditions.
 8. The method ofclaim 1, wherein a specific thrust, defined as a net thrust of theengine divided by a total mass flow through the engine, is between 80Nkg⁻¹s and 95 Nkg⁻¹s at cruise conditions.
 9. The method of claim 1,wherein the fan diameter is between 220 cm and 300 cm.
 10. The method ofclaim 9, wherein the rotational speed of the fan at cruise conditions isin the range of from 1700 rpm to 2500 rpm.
 11. The method of claim 1,wherein the fan diameter is between 320 cm and 380 cm, and therotational speed of the fan at cruise conditions is in the range of from1200 rpm to 2000 rpm.
 12. The method of claim 1, wherein the gearbox hasa planetary configuration and the jet velocity ratio, R_(J), is belowaround 1.0 at cruise conditions.
 13. The method of claim 1, wherein thegearbox has a star configuration and the jet velocity ratio, R_(J), isabove around 1.0 at cruise conditions.
 14. The method of claim 1,wherein a gear ratio of the gearbox is above around 3.4, and the jetvelocity ratio, R_(J), is below around 1.0 at cruise conditions.
 15. Themethod of claim 1, wherein a gear ratio of the gearbox is below around3.4, and the jet velocity ratio, R_(J), is below between around 1.0 and1.3 at cruise conditions.
 16. The method of claim 1, wherein a bypassratio, defined as the ratio of the mass flow rate of the flow throughthe bypass duct to the mass flow rate of the flow through the core atcruise conditions, is in a range of from 10.5 to 15.5.
 17. The method ofclaim 1, wherein the bypass ratio is in a range from 12.5 to 13.5. 18.The method of claim 1, wherein the overall pressure ratio defined as theratio of the stagnation pressure upstream of the fan to the stagnationpressure at an exit of a highest pressure compressor is between 45 and60 at cruise conditions.
 19. The method of claim 1, wherein the fancomprises 16, 18 or 20 fan blades.
 20. The method of claim 1, whereineach fan blade comprises a carbon-fibre or aluminium based body with atitanium leading edge.